1. Field of the Invention
The present invention relates to a measuring instrument of the pressure in a combustor of a gas turbine, rocket engine, etc.
2. Description of the Related Art
A gas turbine or rocket engine is equipped with a combustor. FIG. 7 shows, for example, a sectional view near the combustor of a gas turbine equipped with a combustor 11 to the rotor housing 12. In the drawing, the combustor 11 comprises a fuel nozzle 13, liner 14, and tail tube 15 inside a combustor casing 16. A bypass elbow 17 is connected to the tail tube 15 to which a bypass valve 18 with an adjusting mechanism 19 is attached. Reference number 20 is an air compressor. The air 21 compressed by the compressor 20 flows into the rotor housing 12, proceeds along the outer side of the combustor 11, and is introduced into the combustor 11 from the upstream side of the fuel nozzle 13 as shown by arrows. The fuel injected from the fuel nozzle 13 is burned in the combustor 11, and the combustion gas is introduced to the a gas turbine 22 to drive the turbine rotor.
In the combustor 11 of a gas turbine is produced resonant vibrations of the combustion gas, which are reflected from the components of the combustor 11 such as the liner 14, tail tube 15, and combustor casing 16, etc. The large energy of the resonant vibrations may cause the fatigue failure of the components of the combustor 11 and further the breakage of peripheral parts such as turbine blade, etc. To deal with the problem, the pressure fluctuation of the combustion gas is detected and fuel/air ratio, pilot ratio, the opening of the bypass valve, etc. are adjusted on the basis of the result of the detection.
The pressure in the combustor of a gas turbine or rocket engine is desirable to be measured directly by a pressure sensor located in the combustor, but in that case the pressure sensor directly contacts with the combustion gas of high temperature. A pressure sensor which can withstand such high temperatures is expensive, and also there is a possibility that failed detection of the pressure may occur due to the breakdown of the pressure sensor. If the pressure detection becomes impossible, it causes interference with the operation of the combustor of the gas turbine or rocket engine.
For this reason, a method was proposed to allow the use of an inexpensive pressure sensor, in which the pressure sensor 1 is prevented from being exposed to elevated temperatures by attaching a pressure conduit 2 to the pressure sensor 1 as shown in FIG. 8, inserting the pressure conduit 2 inside the liner 14 through the combustor casing 16 in FIG. 7, thus securing a certain distance between the combustion gas in the liner 14 and the pressure sensor 1.
However, with the pressure conduit 2 inserted into the inside of the liner 14, resonance occurs in the gas column in the pressure conduit 2 in accordance with the eigenvalue of the pressure conduit 2, as a result vibrations not existing in the combustor are created in the pressure conduit 2, the vibration being magnified by the response magnification of the pressure conduit, and the sensor 1 detects a pressure different from that in the liner 14.
To deal with this problem, there was proposed a pressure measuring instrument in Japanese Patent Application Publication No. 6-331146 as shown in FIG. 9, in which a damping tube 9 is provided so that resonance produced in the pressure conduit 2 in accordance with the eigenvalue of the pressure conduit 2 are absorbed and dampened by the damping tube 9 and gas pressure vibration detected is covered with flat amplitudes over the gas vibration frequencies to make it possible for the sensor 1 to accurately measure the pressure.
However, as the damping tube 9 shown in FIG. 9 is located at the position where the temperature is comparatively low (100° C. or below), there used to be such a case that, for example, when the operation is halted and started again after a while, condensation of water occurs in the damping tube 9 and accurate pressure measurement becomes impossible.